The present invention relates generally to gas turbine engines, and, more specifically, to turbine airfoil cooling.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor, and ignited for generating hot combustion gases which flow downstream through turbines for extracting energy therefrom. A high pressure turbine (HPT) immediately follows the combustor and includes a stationary nozzle having a plurality of circumferentially spaced apart stator vanes. The combustion gases are directed by the vanes to engage a row of turbine rotor blades extending outwardly from a supporting rotor disk. Energy is extracted from the gases for rotating the rotor disk, which in turn powers the compressor. A low pressure turbine typically follows a HPT for extracting additional energy from the combustion gases for powering a fan in a typical aircraft engine application.
The nozzle vanes and rotor blades define corresponding airfoils with generally concave, pressure sides and generally convex, suction sides specifically configured for maximizing aerodynamic performance. These turbine airfoils are directly exposed to the hot combustion gases and are cooled by using a portion of air bled from the compressor, and suitably channeled therethrough.
In view of the airfoil configuration of the vanes and blades and their specialized functions in the combustion gas field, the various portions of the airfoils are heated differently by the combustion gases, and correspondingly have different cooling requirements. Since the airfoil leading edges first encounter the combustion gases, they require specialized cooling typically provided by a dedicated cooling cavity therein, with one or more rows of film cooling holes disposed in flow communication therewith.
The mid-chord section of the blade airfoil is typically cooled using a serpentine cooling passage therein with wall turbulators for increasing heat transfer cooling therein. The mid-chord region of the nozzle vane is typically internally cooled with dedicated cooling passages typically including separate impingement baffles for directing jets of cooling air against the inner surfaces of the vane.
The typical profile of a vane or blade airfoil increases in width from the leading edges thereof over the mid-chord regions and tapers in thickness to a thin trailing edge. The thin trailing edges are accordingly difficult to cool in view of the limited space between the pressure and suction sidewalls in which cooling features may be introduced.
A blade trailing edge typically includes a dedicated cooling circuit which is separated from the mid-chord serpentine cooling circuit, and receives a portion of the cooling air at the root of the airfoil for flow radially outwardly along the span of the airfoil for discharge through a row of trailing edge cooling holes. The cooling air turns in the trailing edge cavity to the axial direction for discharge through the trailing edge holes.
Trailing edge cooling effectiveness may be improved by introducing a distributed pattern of small pins formed integrally between the pressure and suction sides of the airfoil. Heat in the trailing edge region of the airfoil is then carried through the pins for extraction by the cooling air channeled therearound which is then discharged out the airfoil trailing edge.
In alternate configurations, the trailing edge cooling cavity may include turbulators in the form of elongate ribs which extend in part from the inner surfaces of the pressure or suction sides of the airfoil and over which the cooling air is channeled. Such turbulator ribs are configured to trip the cooling air flow for promoting turbulence of the air and increased heat transfer cooling. However, as the air flows from rib to rib in turn, it becomes hotter and correspondingly is less effective for cooling downstream turbulators in a series thereof. In this configuration, the turbulators are spaced apart along the span or radial axis of the airfoil between the root and tip thereof.
In yet another configuration, the turbulators may similarly be spaced apart along the span of the trailing edge cavity, but the cooling air flow is delivered thereto from an upstream cavity which directs the cooling air in the axial direction across the ribs for discharge through the trailing edge. In this configuration, the extreme thinness of the trailing edge region of the airfoil and the axially directed cooling air limit the cooling ability of the turbulators.
Accordingly, it is desired to improve airfoil cooling, such as near the trailing edges of turbine blades and vanes.